1. Field of the Invention
The present invention relates generally to a gas turbine engine, and more specifically to a turbine blade with a cooled squealer tip.
2. Description of the Related Art Including Information Disclosed Under 37 CFR 1.97 and 1.98
In a gas turbine engine, a hot gas flow is developed in the combustor from the burning of a fuel with compressed air from the compressor and then passed through a multiple staged turbine to produce mechanical power. In an aero engine, the mechanical power drives the rotor shaft that is connected to a bypass fan. In an industrial gas turbine engine, the rotor shaft is connected to an electric generator that will produce electrical power. In both engines, the engine efficiency can be increased by passing a higher temperature gas into the turbine. However, the turbine inlet temperature is limited to the material properties of the first stage turbine airfoils, these airfoils being the stator vanes and the rotor blades.
Complex internal airfoil cooling passages have been proposed, to provide high levels of airfoil cooling using a minimal amount of cooling air. Higher turbine inlet temperatures are obtainable by providing improved airfoil cooling. Also, since the compressed air used to cool these airfoils is taken from the compressor, the use of a minimal amount of compressor bleed off air for the airfoil cooling will also increase the engine efficiency.
Airfoil cooling is also important in increasing the life of the airfoils. Hot spots can, occur on sections of the airfoils that are not adequately cooled. These hot spots can cause oxidation that will lead to shortened life for the airfoil. Blade tips are especially subject to hot spots since it is nearly impossible to total eliminate the gap between the rotating blade tip and the stationary shroud that forms the gap. Without any gas, blade tip rubbing will occur which leads to other problems. Because of the presence of the tip gap, the hot gas can flow through the gap and expose the blade tip surface to the extreme high temperatures of the gas flow. Therefore, adequate blade tip cooling is also required to reduce hot gas flow leakage and to control metal temperature in order to increase part life.
In the prior art, an airfoil tip edge is cooled by using multiple film cooling holes. FIG. 1 shows a prior art blade tip region with a row of film cooling holes just below the tip edge on the suction side. FIG. 3 shows a row of film cooling holes on the pressure side. FIG. 2 shows a prior art turbine blade. These film cooling holes are fed from the blade internal cavity and exit at various gas side discharge pressures along the blade tip peripheral. As a result of this cooling approach, cooling flow distribution and pressure ratio across these film cooling holes for the pressure side and the suction side film cooling holes are predetermined by the internal cavity pressure. In addition, the blade tip region is also subject to severe secondary flow field which requires a large number of film cooling holes and a large amount of cooling flow to cool the blade tip periphery.